Blade for a rotating bladed disk for an aircraft turbine engine comprising a sealing lip having an optimized non-constant cross section

Abstract

To increase the inertia of a sealing lip of a blade for an aircraft turbine engine, and thus improve the service life of such a sealing lip, the sealing lip is conformed so as to have a trough in the outer surface thereof and a corresponding boss in the inner surface thereof, the trough and the boss being defined based on a connection cross section of the sealing lip to a blade body, and being formed at a distance from a free axial end of the sealing lip.

The invention claimed is: 
     
       1. A blade for a rotating bladed disk for an aircraft turbine engine, comprising:
 a sealing lip extending axially projecting from a body of the blade and having first and second opposite circumferential end surfaces, and one free axial end, 
 wherein a first cross section of the sealing lip at the free axial end is defined by two concentric circle portions extending from the first circumferential end surface to the second circumferential end surface, the first circumferential end surface, and the second circumferential end surface, 
 wherein a second cross section of the sealing lip at a location where the sealing lip is connected to the body of the blade is defined by an internal curved line and by an external curved line, the first circumferential end surface, and the second circumferential end surface, each of the internal and external curved lines comprising: 
 a respective median portion, which is convex in a direction from a head of the blade to a root of the blade; 
 a respective first intermediate portion, which continuously extends from the respective median portion towards the first circumferential end surface of the sealing lip, and which is concave in the direction from the head of the blade to the root of the blade, whereby a respective first point of inflection separates the respective median portion from the respective first intermediate portion; 
 a respective second intermediate portion, which continuously extends from the respective median portion towards the second circumferential end surface of the sealing lip, and which is concave in the direction from the head of the blade to the root of the blade, whereby a respective second point of inflection separates the respective median portion from the respective second intermediate portion; and 
 a respective first end portion and a respective second end portion which respectively continuously extend from the respective first intermediate portion and the respective second intermediate portion, respectively to the first and to the second of the circumferential end surfaces of the sealing lip; 
 each portion of the internal curved line being defined facing the corresponding portion of the external curved line, along the direction from the head of the blade to the root of the blade; and 
 the cross section of the sealing lip evolving in a continuous manner from the second cross section to the free axial end of the sealing lip first cross section. 
 
     
     
       2. The blade according to  claim 1 , wherein the external curved line forms an image of the internal curved line by a homothetic transformation. 
     
     
       3. The blade according to  claim 1 , wherein, for each of the internal and external curved lines, the respective median portion has a respective extremum which is offset, in the direction from the head of the blade to the root of the blade, with respect to the respective circumferential ends of the curved line in question, by a distance which is greater than half a distance separating the respective extrema of the respective median portions of the internal and external curved lines, and which is less than one and a half times the distance separating the respective extrema of the respective median portions of the internal and external curved lines. 
     
     
       4. A rotating bladed disk for an aircraft turbine engine, comprising an annular row of blades according to  claim 1 , distributed about an axis of the rotating bladed disk, and in each whereof the two concentric circle portions, which define the first cross section, have a common center of curvature located on the axis of the rotating bladed disk. 
     
     
       5. The rotating bladed disk according to  claim 4 , wherein, for each of the blades, a homothetic transformation, whereby the external curved line forms an image of the internal curved line, is a proportional transformation wherein the center is located on the axis of the rotating bladed disk. 
     
     
       6. The rotating bladed disk according to  claim 4 , wherein, for each of the blades, the respective first and second end portions are circle portions having a common center of curvature located on the axis of the rotating bladed disk. 
     
     
       7. The rotating bladed disk according to  claim 4 , wherein the respective sealing lips of the blades extend in aerodynamic continuity pairwise. 
     
     
       8. A turbine for an aircraft turbine engine, comprising at least one rotating bladed disk according to  claim 4 . 
     
     
       9. A turbine engine for an aircraft, comprising a high-pressure core and a low-pressure core, wherein the low-pressure core includes a turbine according to  claim 8 . 
     
     
       10. A turbine engine for an aircraft, comprising at least one rotating bladed disk according to  claim 4 .